Gas turbines can be provided with a single combustion chamber, but they can also have what is termed as sequential combustion. In the case of the latter, fuel is burned in a first combustion chamber, and the combustion air is then expanded via a first turbine, a high-pressure turbine. Downstream of the high-pressure turbine, the still hot combustion gases flow through a secondary combustion chamber, in which more fuel is fed in and typically burned in a process involving self-ignition. Arranged downstream of this secondary combustion chamber is a low-pressure turbine, by means of which the combustion gases are expanded, if appropriate followed by a heat recovery system with steam generation.
The transition of the housing from a combustion chamber to a turbine is a critical region here because the temperature and pressure conditions are particularly complex in this region. Typically, the secondary combustion chamber, which is normally designed as an annular combustion chamber, has, as it were, a shell-shaped outer boundary, an outer wall which is composed of a heat-resistant material or is correspondingly coated and which is normally constructed from individual segments. On the opposite, inner side, which is closer to the axis, there is a correspondingly designed inner boundary, an inner wall composed of corresponding materials. The low-pressure turbine, for its part, has a multiplicity of alternately arranged rows of guide vanes and rotor blades. The first row, which is arranged directly downstream of the secondary combustion chamber, is typically a guide vane row exhibiting a considerable twist of the vanes relative to the direction of the principal axis. In this case, the guide vanes are typically designed as segment modules, in which each guide vane has an inner platform on the inside and an outer platform on the outside, and the inner surfaces of these platforms then also form the radially inner and radially outer boundaries of the flow channel for the combustion air.
Accordingly, there is a gap on the radially inner side of the annular flow channel between the inner wall segment of the secondary combustion chamber and the inner platform of the first guide vane row, and a gap on the radially outer side between the outer wall segment of the secondary combustion chamber and the outer platform of the first guide vane row. For reasons of assembly and owing to the different mechanical and thermal loads on the components comprising the secondary combustion chamber and the turbine, this gap must have a certain width and cannot simply be closed or fully bridged. The problem with this gap, which forms a cavity that extends quite a long way radially towards the outside into other structural components of the housing, especially on the radially outer side, is the fact that it is furthermore exposed to complex flow conditions, especially in the region of each guide vane. This is because what is termed a bow wave or a “horse shoe vortex” is formed at the leading edge of the guide vanes, leading to hot combustion air being forced into this cavity in the wall region and penetrating to a corresponding depth into the latter. This can give rise to problems in connection not only with overheating but also with oxidation of the corresponding surfaces.
US 2009/0293488 discloses the possibility of substantially closing this transitional region by means of a very small gap dimension and additionally of providing specific structures which ensure optimum cooling of the wall regions in this region. However, the problem with this approach is that the required clearance between the combustion chamber module and the turbine is not automatically ensured as well, owing to the correspondingly small gap dimension.